Title:
Splice joints for composite aircraft fuselages and other structures
United States Patent 8869403


Abstract:
Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present disclosure includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.



Inventors:
Stulc, Jeffrey F. (Seattle, WA, US)
Chan, Wallace C. (Seattle, WA, US)
Clapp, Brian C. (Seattle, WA, US)
Rolfes, Neal G. (Seattle, WA, US)
Application Number:
13/225057
Publication Date:
10/28/2014
Filing Date:
09/02/2011
Assignee:
The Boeing Company (Chicago, IL, US)
Primary Class:
Other Classes:
29/897, 244/119, 244/120, 244/131, 244/132
International Classes:
B21D53/88
Field of Search:
29/897, 29/897.2, 29/462, 29/464, 29/466, 29/525.01, 244/2, 244/117R, 244/119, 244/120, 244/131, 244/132, 244/133, 52/800.11, 52/801.1
View Patent Images:
US Patent References:
20110073708Composite Barrel Sections for Aircraft Fuselages and Other Structures, and Methods and Systems for Manufacturing Such Barrel SectionsMarch, 2011Biornstad et al.
7716835Methods of manufacturing structural panelsMay, 2010Johnson et al.
7662251Method of manufacturing composite riserFebruary, 2010Salama et al.
7624488One-piece barrel assembly cartDecember, 2009Lum et al.
7556076Tire for two-wheeled vehicle comprising an anti-vibration meansJuly, 2009Prost et al.
20090139641COMPOSITE SECTIONS FOR AIRCRAFT FUSELAGES AND OTHER STRUCTURES, AND METHODS AND SYSTEMS FOR MANUFACTURING SUCH SECTIONSJune, 2009Chapman et al.
7527222Composite barrel sections for aircraft fuselages and other structures, and methods and systems for manufacturing such barrel sectionsMay, 2009Biornstad et al.
20090057487COMPOSITE FABRIC WITH RIGID MEMBER STRUCTUREMarch, 2009Velicki et al.
7503368Composite sections for aircraft fuselages and other structures, and methods and systems for manufacturing such sectionsMarch, 2009Chapman et al.
20080246175Composite Barrel Sections for Aircraft Fuselages and Other Structures, and Methods for Systems for Manufacturing Such Barrel SectionsOctober, 2008Biornstad et al.
7413695Resin infusion mold tool system and vacuum assisted resin transfer molding with subsequent pressure bleedAugust, 2008Thrash et al.
7407556Automated fiber placement using multiple placement heads, replaceable creels, and replaceable placement headsAugust, 2008Oldani et al.
20080111026Splice Joints for Composite Aircraft Fuselages and Other StructuresMay, 2008Stulc et al.
7334782Controlled atmospheric pressure resin infusion processFebruary, 2008Woods et al.
7325771Splice joints for composite aircraft fuselages and other structuresFebruary, 2008Stulc et al.
7289656Systems and methods for determining inconsistency characteristics of a composite structureOctober, 2007Engelbart et al.
7282107Multiple head automated composite laminating machine for the fabrication of large barrel section componentsOctober, 2007Johnson et al.
7278198Mandrel segment loaderOctober, 2007Olson et al.
20070128960Three-dimensionally reinforced multifunctional nanocompositesJune, 2007Ghasemi Nejhad et al.
7236625Systems and method for identifying foreign objects and debris (FOD) and defects during fabrication of a composite structureJune, 2007Engelbart et al.
7228611Method of transferring large uncured composite laminatesJune, 2007Anderson et al.
7195201Adaptable payload enabling architectureMarch, 2007Grether et al.
7193696Systems and methods for using light to indicate defect locations on a composite structureMarch, 2007Engelbart et al.
7171033System and method for identifying defects in a composite structureJanuary, 2007Engelbart et al.
7159822Structural panels for use in aircraft fuselages and other structuresJanuary, 2007Grantham et al.
7141199Method and apparatus for shaping section bar made of composite material and shaped product and I-shaped stringer thereofNovember, 2006Sana et al.
7137182Parallel configuration composite material fabricatorNovember, 2006Nelson
7134629Structural panels for use in aircraft fuselages and other structuresNovember, 2006Johnson et al.
7093797Adaptable payload apparatus and methodsAugust, 2006Grether et al.
7083698Automated composite lay-up to an internal fuselage mandrelAugust, 2006Engwall et al.
7080805Stiffened structures and associated methodsJuly, 2006Prichard et al.
7080441Composite fuselage machine and method of automated composite lay upJuly, 2006Braun
7074474Composite material-stiffened panel and manufacturing method thereofJuly, 2006Toi et al.
20060118244Device for laying tape materials for aerospace applicationsJune, 2006Zaballos et al.
7048024Unidirectional, multi-head fiber placementMay, 2006Clark et al.
7039485Systems and methods enabling automated return to and/or repair of defects with a material placement machineMay, 2006Engelbart et al.
7025305Aircraft panelApril, 2006Folkesson et al.
20050263645Structural panels for use in aircraft fuselages and other structuresDecember, 2005Johnson et al.
20050163965Molding process and apparatus for producing unified composite structuresJuly, 2005Velicki et al.
6910043Compression of nodes in a trie structureJune, 2005Iivonen et al.
6896841Molding process and apparatus for producing unified composite structuresMay, 2005Velicki et al.
6871684System for identifying defects in a composite structureMarch, 2005Engelbart et al.
6860957Automatic prepreg laminating method and apparatus for carrying out the sameMarch, 2005Sana et al.
6840750Resin infusion mold tool system and vacuum assisted resin transfer molding with subsequent pressure bleedJanuary, 2005Thrash et al.
6817574Structural element for an aircraft, especially an aircraft doorNovember, 2004Solanille et al.
6814822Composite material collation machine and associated method for high rate collation of composite materialsNovember, 2004Holmes et al.
6802931Method for producing composite structureOctober, 2004Fujihira
6799619Composite material collation machine and associated method for high rate collation of composite materialsOctober, 2004Holmes et al.
6786452Wing structure of airplaneSeptember, 2004Yamashita et al.
6779707Friction stir welding as a rivet replacement technologyAugust, 2004Dracup et al.
6766984Stiffeners for aircraft structural panelsJuly, 2004Ochoa
6743504Co-cured composite structures and method of making themJune, 2004Allen et al.
6730184Method for producing fiber-reinforced composite semi-hardened product having joggle, and method for producing preformed structure using sameMay, 2004Kondo et al.
6709538Method of making a laminated composite radius fillerMarch, 2004George et al.
6702911Composite material-stiffened panel and manufacturing method thereofMarch, 2004Toi et al.
20040035979Integrally stiffened axial load carrying skin panels for primary aircraft structure and closed loop manufacturing methods for making the sameFebruary, 2004McCoskey, Jr. et al.
6692681Method and apparatus for manufacturing composite structuresFebruary, 2004Lunde
6663737Core-crush resistant fabric and prepreg for fiber reinforced composite sandwich structuresDecember, 2003Hsiao et al.
6648273Light weight and high strength fuselageNovember, 2003Anast
20030190455Textile joint reinforcement and associated methodOctober, 2003Burgess et al.
6622974Geometric morphing wing with expandable sparsSeptember, 2003Dockter et al.
6620484Variable density stitched-composite structural elements for energy absorptionSeptember, 2003Bolukbasi et al.
6613258Method for making parts in composite material with thermoplastic matrixSeptember, 2003Maison et al.
6589618Resin transfer molding processJuly, 2003Cundiff et al.
6561478Resin transfer moldMay, 2003Cundiff et al.
6560843Accurate positioning using a seal pin for RTM mold diesMay, 2003Cundiff et al.
6547769Catheter apparatus with weeping tip and method of useApril, 2003VanTassel et al.
6511570Method for producing body structure of fiber-reinforced composite, and body structure produced therebyJanuary, 2003Matsui
6510961Integrally-reinforced braided tubular structure and method of producing the sameJanuary, 2003Head et al.
6508909Process for manufacturing pre-cured parts of composite material with green-applied stiffenersJanuary, 2003Cerezo Pancorbo et al.
6480271Traversing laser locating systemNovember, 2002Cloud et al.
6451152Method for heating and controlling temperature of composite material during automated placementSeptember, 2002Holmes et al.
6431837Stitched composite fan bladeAugust, 2002Velicki
6415581Corrugated stiffening memberJuly, 2002Shipman et al.
6390169Conformable compaction apparatus for use with a fiber placement machineMay, 2002Johnson
6374750Structural panel systemApril, 2002Early
6364250Shell component for an aircraft fuselage and method of manufacturing the sameApril, 2002Brinck et al.
6319447Resin transfer molding processNovember, 2001Cundiff et al.
6231941Radius fillers for a resin transfer molding processMay, 2001Cundiff et al.
6205239System and method for circuit repairMarch, 2001Lin et al.
6198983Table-driven software architecture for a stitching systemMarch, 2001Thrash et al.
6190484Monolithic composite wing manufacturing processFebruary, 2001Appa
6187411Stitch-reinforced sandwich panel and method of making sameFebruary, 2001Palmer
6168358Hybrid lay-up toolJanuary, 2001Engwall et al.
6155450Composite shell shaped as a body of revolutionDecember, 2000Vasiliev et al.
6136237Method of fabricating a fiber-reinforced ceramic matrix composite partOctober, 2000Straub et al.
6129031Robotic stitching apparatus and end effector thereforOctober, 2000Sarh et al.
6128545Automated apparatus and method of generating native code for a stitching machineOctober, 2000Miller
6114050Titanium-polymer hybrid laminatesSeptember, 2000Westre et al.
6114012Rib of composite material and method of forming the sameSeptember, 2000Amaoka et al.
6112792Fiber placement mid-span redirectSeptember, 2000Barr et al.
6099906Immersion process for impregnation of resin into preformsAugust, 2000Palmer et al.
6086696Method of forming a seamless, cylindrical, thermoplastic structure with a multiple compaction roller winderJuly, 2000Gallagher
6074716Weavable metal matrix impregnated tow composite materialJune, 2000Tsotsis
6070831Aircraft for passenger and/or cargo transportJune, 2000Vassiliev et al.
6051089Reinforcing member for composite workpieces and associated methods2000-04-18Palmer et al.
6045651Hand assisted lamination system2000-04-04Kline et al.
6013341Carrying (bearing) pipe-casing made of composite materials, the method and the setting (straightening device) for its manufacturing2000-01-11Medvedev et al.
6012883Hybrid lay-up tool2000-01-11Engwall et al.
6003812Airplane fuselage panel1999-12-21Micale et al.
5979531Bi-directional fiber placement head1999-11-09Barr et al.
5963660Method and apparatus for detecting and measuring laps and gaps in composite materials1999-10-05Koontz et al.
5954917Automated material delivery system1999-09-21Jackson et al.
5953231Automated quality control for stitching of textile articles1999-09-14Miller et al.
5951800Fiber/metal laminate splice1999-09-14Pettit
5931107Advanced stitching head for making stitches in a textile article having variable thickness1999-08-03Thrash et al.
5915317Automated gantry-type stitching system1999-06-29Thrash et al.
5902535Resin film infusion mold tooling and molding method1999-05-11Burgess et al.
5893534Structural apparatus and design to prevent oil can movement of webs in aircraft pressure bulkheads1999-04-13Watanabe
5871117Tubular load-bearing composite structure1999-02-16Protasov et al.
5814386Composite shell formed as a body of rotation, and method and mandrel for making same1998-09-29Vasiliev et al.
5809805Warp/knit reinforced structural fabric1998-09-22Palmer et al.
5804276Composite structure adapted for controlled structural deformation1998-09-08Jacobs et al.
5765329Roof construction of corrugated sheets1998-06-16Huang
5746553Dual purpose lay-up tool1998-05-05Engwall
5700337Fabrication method for composite structure adapted for controlled structural deformation1997-12-23Jacobs et al.
5683646Fabrication of large hollow composite structure with precisely defined outer surface1997-11-04Reiling, Jr.
5651600Method for controlling projection of optical layup template utilizing cooperative targets1997-07-29Dorsey-Palmateer
5622733Tooling for the fabrication of composite hollow crown-stiffened skins and panels1997-04-22Asher
5619837Corrugated panel structure1997-04-15DiSanto
5562788Composite material laser flaw detection1996-10-08Kitson et al.
5540126Automatic lay-up machine for composite fiber tape1996-07-30Piramoon
5518208Optimum aircraft body frame to body skin shear tie installation pattern for body skin/stringer circumferential splices1996-05-21Roseburg
5450147Method for controlling projection of optical layup template utilizing cooperative targets1995-09-12Dorsey-Palmateer
5439549Double edged pressure sensitive folded tape application apparatus1995-08-08Fryc et al.
5429326Spliced laminate for aircraft fuselage1995-07-04Garesche et al.
5399406Paneling material and composite panel using the same1995-03-21Matsuo et al.
5384959Method of making a SPF/DB hollow core fan blade1995-01-31Velicki
53376473 dimensional braiding apparatus1994-08-16Roberts et al.
5297760Aircraft skin lap splice1994-03-29Hart-Smith
5281388Resin impregnation process for producing a resin-fiber composite1994-01-25Palmer et al.
5262220High strength structure assembly and method of making the same1993-11-16Spriggs et al.
5251849Strain reduced airplane skin1993-10-12Torres
5242523Caul and method for bonding and curing intricate composite structures1993-09-07Willden et al.
5240376SPF/DB hollow core fan blade1993-08-31Velicki
5223067Method of fabricating aircraft fuselage structure1993-06-29Hamamoto et al.
5148588Process for the manufacture of a fabric-covered visor1992-09-22Prillard
5086997Structural joint and a method for joining in reinforced thermoplastic fabrication1992-02-11Glass
5058497Compliant pressure roller1991-10-22Bishop et al.
5024399Frame made of a composite material, especially for the fuselage of an aircraft, and its method of production1991-06-18Barquet et al.
4966802Composites made of fiber reinforced resin elements joined by adhesive1990-10-30Hertzberg
4959220Antiseptic-containing alginate impression material1990-09-25Yamamoto et al.
4942013Vacuum resin impregnation process1990-07-17Palmer et al.
4941182Vision system and method for automated painting equipment1990-07-10Patel
4877471Method and apparatus for delivering a resin-impregnated, multifilament band1989-10-31McCowin et al.
4830298Self-centering sheave for filaments1989-05-16Van Blunk
4828202Method and apparatus for wideband vibration damping of reinforced skin structures1989-05-09Jacobs et al.
4811540Fiber reinforced shell structure of synthetic material1989-03-14Kallies et al.
4790898Method and apparatus for fiber lamination1988-12-13Woods
4780262Method for making composite structures1988-10-25VonVolkli
4760444Machine visual inspection device and method1988-07-26Nielson et al.
4736566Modular fabrication panel system1988-04-12Krotsch
4715560Composite cruciform structure for joining intersecting structural members of an airframe and the like1987-12-29Loyek
4699683Multiroving fiber laminator1987-10-13McCowin
4693678Male layup-female molding system for fabricating reinforced composite structures1987-09-15Von Volkli
4631221Sheet-like sandwich molding1986-12-23Disselbeck et al.
4622091Resin film infusion process and apparatus1986-11-11Letterman
4615935Glass fiber reinforced ceramic preform and method of casting it1986-10-07Bendig et al.
4608220Method of forming composite material articles1986-08-26Caldwell et al.
4571355Fiber reinforced resin composites formed of basic ply blankets1986-02-18Elrod
4548859Breather material and method of coating fabric with silicone rubber1985-10-22Kline et al.
4548017Building panel1985-10-22Blando
4546717Sewing machine differential feed1985-10-15Marchesi
4542055Three-dimensional fabric painting surfaces1985-09-17Fitzsimmons
4492607Method for producing integrally stiffened fiber reinforced plastic panels1985-01-08Halcomb
4490958Sheet metal beam1985-01-01Lowe
4463044Composite panel of varied thickness1984-07-31McKinney
4448838Graphite fiber reinforced laminate structure capable of withstanding lightning strikes1984-05-15McClenahan et al.
4410577Woven layered cloth reinforcement for structural components1983-10-18Palmer et al.
4331723Advanced composite1982-05-25Hamm
4331495Method of fabricating a reinforced composite structure1982-05-25Lackman et al.
4311661Resin impregnation process1982-01-19Palmer
4310132Fuselage structure using advanced technology fiber reinforced composites1982-01-12Robinson et al.
4256790Reinforced composite structure and method of fabrication thereof1981-03-17Lackman et al.
4186535Shear load resistant structure1980-02-05Morton
4086378Stiffened composite structural member and method of fabrication1978-04-25Kam et al.
4064534System for monitoring the production of items which are initially difficult to physically inspect1977-12-20Chen et al.
3995080Filament reinforced structural shapes1976-11-30Cogburn et al.
3976269Intrinsically tuned structural panel1976-08-24Gupta
3974313Projectile energy absorbing protective barrier1976-08-10James
3879245METHOD OF MAKING COMPOSITE CORED STRUCTURES1975-04-22Fetherston et al.
3603096APPARATUS FOR INSTALLING A REINFORCED VESSEL IN AN UNDERGROUND CAVITY1971-09-07Wells
3507634COMPOSITE METAL STRUCTURE1970-04-21O'Driscoll
3490983FIBER REINFORCED STRUCTURES AND METHODS OF MAKING THE SAME1970-01-20Lee
3452501SNAP LOCKING STRUCTURAL DEVICE1969-07-01Sickler et al.
3306797Method and apparatus for making elongated articles of fiber reinforced resin material1967-02-28Boggs
3271917Reinforced plastic constructions1966-09-13Rubenstein
3071217Vibration damping in sheet metal structures1963-01-01Gould
2992711Reinforcing means for attaching structural members to lightweight corrugated panels1961-07-18Mitchell et al.
2387219Aircraft structure1945-10-16Wallis
2367750Aircraft construction1945-01-23Berkow et al.
2292372Structural element1942-08-11Gerlach et al.
1976257Laminated body and method of making same1934-10-09Harper
0002004N/A1841-03-12Harris et al.



Foreign References:
DE03040838May, 1982
DE3331494March, 1985Multi-web laying machine
EP0319797June, 1989Method and apparatus for measuring defect density and defect distribution
EP0833146April, 1998Method and apparatus for detecting and measuring laps and gaps in composite materials
EP1149687October, 2001Method for producing body structure of fiber-reinforced composite, and body structure produced thereby
GB2224000April, 1990AIRCRAFT FUSELAGE OR OTHER PRESSURE VESSEL.
JP61169394July, 1986
WO/1998/032589January, 1998METHOD AND APPARATUS FOR MANUFACTURING COMPOSITE STRUCTURES
WO/2003/035380May, 2003METHOD FOR MAKING EXTRUDED PROFILES HAVING A SPECIFIC SURFACE STATE MADE OF FIBER-REINFORCED SYNTHETIC RESINS AND MACHINE THEREFOR
Other References:
“Rocky Mountain Composites,” 1 page, accessed Feb. 28, 2004, http://www.rockymountaincomposites.com/wind-sys.htm.
“Beechcraft's Composite Challenge,” 2 page, accessed Mar. 1, 2004, http://www.aerotalk.com/Beech.cfm.
“Business Aviation,” Jun. 7, 2002, 2 pages, accessed Mar. 1, 2004, http://www..aviationnow.com/avnow/news/channel—busay.jsp?view=story&id=news/btoyo0607.xml.
“Raytheon Aircraft Orders Four More Fiber Cincinnnati Fiber Placement Systems for Industry's First Composite Fuselage Business Jets,” Jul. 20, 2000, 2 pages.
“Raytheon's New Quiet Jets,” Vibro-Acoustic Sciences Newsletter, vol. 4, No. 2, Mar. 2000, 2 pages.
“The Barrelful of Experience,” Intervia, May 1992, 2 pp.
“CNC fiber placement used to create an all-composite fuselage,” Aerospace Engineering Online, 3 pp., accessed Aug. 31, 2006, http://www.sae.org/aeromag/techinnovations/1298t08.htm.
“Raytheon Aircraft's Hawker Horizon Reaches Fulelage Milestone,” press release, 3 pp., accessed Aug. 31, 2006, http://www.beechcraft.de/Presse/2000/100900b.htm.
Evans, “Fiber Placement,” Manufacturing Processes, ASM Handbook, vol. 21, 2001, pp. 477-479.
“Premier: Features Lighter, Stronger All-Composite Fuselage,” WolfTracks Company Magazines, vol. 4, No. 1, 1998, 3 pp.
Grimshaw, “Automated Tape Laying,” ASM Handbook vol. 21, Composites (ASM International), 2001, pp. 480-485.
Fiedler et al., “Tango Composite Fuselage Platform,” SAMPE Journal, vol. 39, No. 1, Jan./Feb. 2003, 8 pp.
Sharp et al., “Material Selection/Fabrication issues for Thermoplastic Fiber Placement,” Journal of Thermoplastic Composite Materials, vol. 8, Jan. 1995, pp. 2-14.
Ando et al., “Growing Carbon Nanotubes,” Materialsl Today, Oct. 2004, pp. 22-29.
“Growing Carbon Nanotubes Aligned With Patterns,” NASA's Jet Propulsion Laboratory, Pasadena, CA, 4 pp., accessed Mar. 21, 2007, http:www.nasatech.com/Briefs/Oct02/NPO30205.html.
“The Longest Carbon Nanotubes You Have Ever Seen,” Space Mart, May 14, 2007, 3 pages.
Velicki et al., “Damage Arrest Design Approach Using Stitched Composites,” 2nd Aircraft Structural Design Conference, 2010, pp. 1-9.
USPTO Notice of allowance for U.S. Appl. No. 10/949,848 dated Aug. 14, 2007.
PCT International Search Report and Written Opinion for PCT/US2005/032737; Applicant: The Boeing Company; dated Dec. 19, 2006; 16 pgs; European Patent Office.
Grimshaw et al. ,“Advanced Technology Tape Laying for Affordable Manufacturing of Large Composite Structures”, (11 pgs); http://www.cinmach.com/tech/pdf/TapeLayingGrimshaw.pdf.
International Search Report and Written Opinion for PCT/US2004/039905; Applicant: The Boeing Company; May 25, 2005; 10 pgs.
Zhang, “Angewandte Sensorik” CH 4, Sensoren in der Robotik, Nov. 11, 2003, pp. 76-113; XP002327793; URL:http://tech-www.informatik.uni-hamburg.de/lehre/ws2003/veriesungen/angewandte—sensorik/verlesung—03.pdf, accessed Apr. 2004.
USPTO Office action for U.S. Appl. No. 10/949,848 dated Sep. 28, 2006.
USPTO Notice of allowance for U.S. Appl. No. 12/016,258 dated Jul. 19, 2011.
USPTO Notice of allowance for U.S. Appl. No. 12/016,258 dated Mar. 16, 2011.
Garcia et al., “Hybrid Carbon Nanotube-Composite Architectures”, MTL Annual Research Report, Sep. 2006, 1 pg.
“The Longest Carbon Nanotubes You Have Ever Seen”, http://www.spacemart.com/reports/The—Longest—Carbon—Nanotubes—You—Have—Ever—Seen—999.html, May 14, 2007, 1 page.
“The Wondrous World of Carbon Nanotubes”, Eindhoven University of Technology, Feb. 2003, http://students.chem.tue.nl/ifp03/synthesis.html, accessed Mar. 21, 2007, 96 pgs.
Primary Examiner:
WALTERS, RYAN J
Attorney, Agent or Firm:
Boeing Management Company - IP Management (BOEING MANAGEMENT COMPANY 5301 BOLSA AVENUE MAIL CODE H011-B171, Huntington Beach, CA, 92647, US)
Parent Case Data:
This application is a divisional of U.S. application Ser. No. 12/016,258, filed Jan. 18, 2008, U.S. Pat. No. 8,061,035, which is a divisional of U.S. application Ser. No. 10/949,848, filed Sep. 23, 2004. U.S. Pat. No. 7,325,771, both of which are incorporated herein by reference.
Claims:
What is claimed is:

1. A method for manufacturing a shell structure, the method comprising: attaching at least a first stiffener to a first skin having a first edge cut-out portion; attaching at least a second stiffener to a second skin having a second edge cut-out portion; positioning the first skin in edgewise alignment with the second skin whereby the first edge cut-out portion is at least approximately aligned with the second edge cut-out portion; and attaching a first end of a fitting to the first stiffener and the first skin and a second end of the fitting to the second stiffener and the second skin attaching a strap to a first edge region of the first skin and a second edge region of the second skin to splice the first and second skins together, wherein the strap includes an aperture at least approximately aligned with the first and second edge cut-out portions of the first and second skins.

2. The method of claim 1, wherein attaching the fitting includes sandwiching a portion of the strap between the fitting and the first edge region of the first skin and the second edge region of the second skin.

Description:

BACKGROUND INFORMATION

1. Field

The following disclosure relates generally to shell structures and, more particularly, to splice joints for joining composite fuselage sections and other shell structures together.

2. Background

The primary structural elements of large passenger jets and other large aircraft are typically made from metal. Fuselage shells for such aircraft, for example, are typically manufactured from high-strength aluminum alloys or similar metals. In an effort to increase performance, however, many aircraft manufacturers are turning to fiber-reinforced resin materials (i.e., “composite” materials) that have relatively high strength-to-weight ratios. Conventional composite materials typically include glass, carbon, or polyaramide fibers in a matrix of epoxy or another type of resin. The use of such materials for primary structures has mostly been limited to smaller aircraft, such as fighter aircraft, high-performance private aircraft, and business jets.

One known method for manufacturing business jet airframes with composite materials is employed by the Raytheon Aircraft Company of Wichita, Kans., to manufacture the Premier I and Hawker Horizon business jets. This method involves wrapping carbon fibers around a rotating mandrel with an automated fiber placement system. The mandrel provides the basic shape of a longitudinal fuselage section. The carbon fibers are preimpregnated with a thermoset epoxy resin, and are applied over the rotating mandrel in multiple plies to form an interior skin of the fuselage section. The interior skin is then covered with a layer of honeycomb core. The fiber placement system then applies additional plies of preimpregnated carbon fibers over the honeycomb core to form an exterior skin that results in a composite sandwich structure.

The Premier I fuselage includes two 360-degree sections formed in the foregoing manner. The Hawker Horizon fuselage includes three such sections formed in this manner. The two 70-inch diameter sections of the Premier I fuselage are riveted and then bonded together at a circumferential splice joint to form the complete fuselage structure. The much larger Hawker Horizon fuselage, with an 84-inch diameter, uses aluminum splice plates at two circumferential joints to join the three fuselage sections together into a complete structure.

To precisely install the aluminum splice plates on the Hawker Horizon fuselage, Raytheon created a special, automated splice machine. This machine aligns the three fuselage sections using a computer-aided laser alignment system, and then drills attachment holes through the aluminum splice plates and the underlying sandwich structure. The machine then probes each hole for size quality and records statistical process control data on each hole. The drill heads also apply sealant and install hi-shear fasteners in approximately 1,800 places along each of the splice joints. (See Raytheon Aircraft news release at http://www.beechcraft.de/presse/2000/100900b.htm entitled “RAYTHEON AIRCRAFT'S HAWKER HORIZON REACHES FUSELAGE MILESTONE,” Oct. 9, 2000).

SUMMARY

The present disclosure is directed generally toward structures and methods for joining composite fuselage sections and other panel assemblies together. A shell structure configured in accordance with one aspect of the invention includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can further include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to the second stiffener and the second composite skin, to join the first panel portion to the second panel portion.

A method for manufacturing a shell structure in accordance with another aspect of the invention includes attaching at least a first stiffener to a first composite skin, and attaching at least a second stiffener to a second composite skin. The method can further include positioning the first composite skin in edgewise alignment with the second composite skin, attaching a first end of a fitting to the first stiffener and the first composite skin, and attaching a second end of the fitting to the second stiffener and the second composite skin. In one embodiment, the method can additionally include attaching a strap to a first edge region of the first composite skin and an adjacent second edge region of the second composite skin to splice the first and second composite skins together before the fitting is attached.

The features, functions, and advantages can be achieved independently in various embodiments of the present disclosure or may be combined in yet other embodiments in which further details can be seen with reference to the following description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the advantageous embodiments are set forth in the appended claims. The advantageous embodiments, however, as well as a preferred mode of use, further objectives and advantages thereof, will best be understood by reference to the following detailed description of an advantageous embodiment of the present disclosure when read in conjunction with the accompanying drawings, wherein:

FIG. 1 is an isometric view of an aircraft having a composite fuselage configured in accordance with an embodiment of the invention.

FIGS. 2A-2C together illustrate a method of joining a first fuselage barrel section to a second fuselage barrel section in accordance with an embodiment of the invention.

FIGS. 3A-3C together illustrate a method of joining the first fuselage barrel section to the second fuselage barrel section in the vicinity of a window cutout, in accordance with another embodiment of the invention.

FIG. 4 is a cross-sectional end view of the splice joint of FIG. 2C taken substantially along line 4-4 in FIG. 2C.

DETAILED DESCRIPTION

The following disclosure describes structures and methods for joining composite fuselage sections and other panel assemblies together. Certain details are set forth in the following description and in FIGS. 1-3C to provide a thorough understanding of various embodiments of the invention. Other details describing well-known structures and systems often associated with composite parts and related assembly techniques are not set forth in the following disclosure to avoid unnecessarily obscuring the description of the various embodiments of the invention.

Many of the details, dimensions, angles, and other features shown in the Figures are merely illustrative of particular embodiments of the invention. Accordingly, other embodiments can have other details, dimensions, angles, and features without departing from the spirit or scope of the present invention. In addition, further embodiments of the invention can be practiced without several of the details described below.

In the Figures, identical reference numbers identify identical or at least generally similar elements. To facilitate the discussion of any particular element, the most significant digit or digits of any reference number refer to the Figure in which that element is first introduced. For example, element 106 is first introduced and discussed with reference to FIG. 1.

FIG. 1 is an isometric view of an aircraft 100 having a composite fuselage 102 configured in accordance with an embodiment of the invention. In one aspect of this embodiment, the fuselage 102 includes a plurality of composite barrel sections 104 (identified individually as barrel sections 104a-e) joined together by a plurality of corresponding splice joints 106 (identified individually as splice joints 106a-f). Each of the barrel sections 104 includes a composite skin 112 (identified individually as composite skins 112a-112e) extending 360 degrees around a longitudinal axis 108. In the illustrated embodiment, each of the composite skins 112 can have a cross-sectional width of at least about 10 feet, such as about 15 feet to about 35 feet. In one embodiment, for example, the composite skins 112 can have a cross-sectional width of about 18 feet. Throughout this disclosure, the term “barrel section” is used for convenience to refer to any shell structure extending 360 degrees around an axis. Accordingly, the term is not limited to cylindrical structures or structures having barrel shapes, but can include structures having circular, elliptical, oval, egg-shaped, rectilinear, tapered, or other cross-sectional shapes. In addition, in one embodiment, the barrel sections 104 can be “one-piece” barrel sections in which the composite skins 112 are “one-piece” skins extending continuously for 360 degrees around the axis. In other embodiments, however, the skins 112 can be formed from two or more skin segments spliced or otherwise joined together to form the full 360-degree barrel section.

The fuselage 102 can further include a passenger cabin 103 configured to hold a plurality of passenger seats 105 ranging in number from about 50 to about 700 seats. For example, in the illustrated embodiment, the passenger cabin 103 can hold from about 150 to about 600 passenger seats 105. In other embodiments, the passenger cabin 103 can be configured to hold more or fewer passenger seats without departing from the spirit or scope of the present disclosure. Each of the barrel sections 104 can include a plurality of window cutouts 140 to provide the passengers seated in the passenger cabin 103 with views out of the aircraft 100.

FIGS. 2A-2C together illustrate a method of joining the first barrel section 104a to the second barrel section 104b in accordance with an embodiment of the invention. Referring first to FIG. 2A, this view is a partially exploded, enlarged isometric view looking outwardly at a portion of the second splice joint 106b from within the fuselage 102 (FIG. 1). The portion of the first barrel section 104a illustrated in FIG. 2A includes a first panel portion 210a. The portion of the second barrel section 104b illustrated in FIG. 2A includes a second panel portion 210b positioned in edgewise alignment with the first panel portion 210a. In one embodiment, the panel portions 210 can be at least generally similar in structure and function to the panel assemblies described in detail in co-pending U.S. patent application Ser. No. 10/851,381, filed May 20, 2004, and Ser. No. 10/853,075, filed May 25, 2004, both of which are incorporated herein in their entireties by reference. For example, the first panel portion 210a can include a plurality of stiffeners 214 (identified individually as stiffeners 214a-214e) attached to the first skin 112a. Each of the stiffeners 214 can include a raised portion 224 projecting away from the first skin 112a, and a plurality of flange portions (identified individually as first flange portions 226a and second flange portions 226b) attached directly to the first skin 112a. In the illustrated embodiment, the stiffeners 214 have hat-shaped cross-sections. In other embodiments, however, the stiffeners 214 can have other cross-sectional shapes, including “L” shapes, “C” shapes, inverted “T” shapes, “I” shapes, etc. In yet other embodiments, the panel portions 210 can include other features, including those disclosed in co-pending U.S. patent application Ser. No. 10/819,084, filed Apr. 6, 2004, and incorporated herein in its entirety by reference.

The stiffeners 214 can be positioned on the first skin 112a so that the first flange portions 226a of one stiffener 214 are aligned with the corresponding second flange portions 226b of an adjacent stiffener 214. By aligning the flange portions 226 in the foregoing manner, the flange portions 226 can form a plurality of at least approximately continuous support surfaces 228 (identified individually as support surfaces 228a and 228b) extending between the raised portions 224 of the stiffeners 214.

The first panel portion 210a can further include part of a support member or frame 216a. In the illustrated embodiment, the frame 216a is a two-piece frame that includes a first frame section 218 and a second frame section 219. The first frame section 218 can be attached directly to the support surfaces 228 as described in detail in U.S. patent application Ser. No. 10/851,381. In other embodiments, the first frame section 218 can be attached to the first panel portion 210a using other methods. In still further embodiments, the first panel portion 210a can include parts of other frames composed of more or fewer frame sections. Alternatively, the frame 216a can be omitted.

The second panel portion 210b can be at least generally similar in structure and function to the first panel portion 210a described above. Accordingly, the second panel portion 210b can include a plurality of stiffeners 214 (identified individually as stiffeners 214f-j) attached to the second skin 112b. The second panel portion 210b can further include a second frame 216b that is attached to flange portions of the stiffeners 214 in the manner described above for the first panel portion 210a.

Referring next to FIG. 2B, an elongate strap 220 is attached to a first edge region 213a of the first skin 112a and an adjacent second edge region 213b of the second skin 112b to splice the first skin 112a to the second skin 112b. The strap 220 is attached to the inner side of the respective skins 112 to maintain a smooth, aerodynamic surface on the exterior of the fuselage 102 (FIG. 1). In one embodiment, the strap 220 can include composite materials, such as graphite-epoxy or similar material. In other embodiments, the strap 220 can include other materials, including metallic materials such as aluminum, titanium, steel, etc. The strap 220 can be attached to the skins 112 with a plurality of fasteners 221 extending through the strap 220 and the skins 112. In other embodiments, the strap 220 can be bonded to the skins 112, or bonded and fastened to the skins 112. Further, in embodiment, the strap 220 can extend continuously, or at least approximately continuously, around the splice joint 106b. In other embodiments, the strap 220 can be segmented around the splice joint 106b. For example, in one embodiment, the splice joint 106b can include six segments of the strap 220. In other embodiments, more (e.g., eight) or less segments of the strap 220 can be used.

In the illustrated embodiment, the strap 220 can be at least approximately as thick as the skins 112, but thicker than the adjacent flange portions 226 of the stiffeners 214. To avoid a step between adjacent surfaces, shim pads or fillers 222 (identified individually as first fillers 222a and second fillers 222b) are positioned on the flange portions 226 adjacent to the strap 220. In one embodiment, the fillers 222 can include composite materials, including graphite-epoxy or similar materials. In other embodiments, the fillers 222 can include aluminum and other metals. In yet other embodiments, the strap 220, the skins 112, and/or the flange portions 226 can have other relative thicknesses and/or the fillers 222 can be omitted.

Referring next to FIG. 2C, a plurality of fittings 230 are positioned on the strap 220 and extend across the splice joint 106b between the stiffeners 214. A first end portion 232a of each fitting 230 overlays the corresponding first filler 222a and the flange portions 226 of the adjacent stiffeners 214. Similarly, a second end portion 232b of each fitting 230 overlays the corresponding second filler 222b and the flange portions 226 of the adjacent stiffeners 214. In the illustrated embodiment, each of the fittings 230 has a channel or “U-shaped” cross section that includes a base portion 234, a first upstanding edge portion 236a positioned toward a first side of the base portion 234, and a second upstanding edge portion 236b positioned toward a second side of the base portion 234. In other embodiments, the fittings 230 can have other cross-sectional shapes, including “C” shapes, “L” shapes, inverted “Pi” shapes, and flat shapes, to name a few. A plurality of fasteners 238 extending through the fittings 230 and the underlying structures (i.e., the fillers 222, the flange portions 226, the strap 220, and the skins 112) attach the fittings 230 to the underlying structures to form a structural load path across the splice joint 106b.

The fittings 230, the stiffeners 214, the strap 220, and the skins 112 can include composite materials, including graphite-epoxy and/or other suitable composite materials. For example, in one embodiment, the skins 112 can be manufactured with toughened epoxy resin and carbon fibers, e.g., intermediate carbon fibers from Toray Composites America, Inc. of 19002 50th Avenue East, Tacoma, Wash. 98446. In this embodiment, the skins 112 can include fiber tape pre-impregnated with resin (i.e., “prepreg”) and outer plies of prepreg fabric. In another embodiment, the strap 220 and the fittings 230 can also be manufactured from epoxy resin and carbon fibers. The skins 112, the strap 220, and the fittings 230 can have quasi-isotropic lay-ups, i.e., lay-ups having an equal (or approximately equal) number of plies with 0, +45, −45, and 90 degree orientations. The stiffeners 214 can have axial-dominated fiber orientations. In other embodiments, the skins 112, the strap 220, the fittings 230, and the stiffeners 214 can have other fiber orientations.

One advantage of using composite materials instead of metals is that the fittings 230 and the underlying structures (e.g., the skins 112 and the stiffeners 214) will have at least generally similar coefficients of thermal expansion. As a result, temperature fluctuations experienced during operation of the aircraft 100 (FIG. 1) will not cause disparate thermal expansion between the fittings 230 and the underlying structures, and hence will not induce significant stresses in the splice joint 106b. In other embodiments, however, the fittings 230 can include metal materials such as aluminum, titanium, steel, etc. The use of metals may be appropriate in those situations in which the aircraft is not expected to experience wide temperature fluctuations during operation.

In addition to composites and metal materials, in yet other embodiments, the skins 112, the strap 220, the fittings 230, and the stiffeners 214, and combinations thereof, can include other materials, including hybrid materials such as fiber/metal laminates. Such laminates include fiberglass/aluminum laminates and titanium reinforced graphite laminates (Ti/Gr). One hybrid laminate that includes alternating layers of aluminum and fiberglass is referred to as “GLARE™.” This laminate may offer better fatigue properties than conventional aluminum. A Ti/Gr laminate may offer weight advantages over conventional aluminum or graphite-epoxy, but this laminate may also be more expensive.

One feature of the splice joint 106b illustrated in FIG. 2C is that the fittings 230 overlap the strap 220. One advantage of this feature is that it provides a fail-safe, redundant load path in the unlikely event that a crack or other structural flaw propagates through a portion of the strap 220. In such an event, the fittings 230 alone can carry the structural load across the splice joint 106b. In addition, the fittings 230 also provide a redundant load path across the splice joint 106b from where the stiffeners 214 terminate. Further, if a segmented strap 220 is used, then the fittings 230 can also be used as splice plates for adjacent strap segments. Another feature of the splice joint 106b is that the ends of the stiffeners 214 are left open. One advantage of this feature is that it enables moisture caused by condensation and other sources to escape the stiffeners 214 for sufficient drainage.

One feature of the fittings 230 of the illustrated embodiment are the first and second upstanding edge portions 236a and 236b. The upstanding edge portions 236 can add stiffness to the fittings 230, and can be positioned proximate to the raised portions 224 of the stiffeners 214. One advantage of this configuration is that it can increase the stability of the splice joint 106b, especially under compression loads.

Yet another feature of the illustrated embodiment is that the raised portions 224 of opposing stiffeners 214 are not spliced together across the splice joint 106b. One advantage of this feature is that it makes the fittings 230 relatively easy to install because the raised portions 224 do not have to be in perfect alignment. While the raised portions 224 could be spliced together in other embodiments, doing so would most likely add time and cost to manufacturing of the splice joint because of the various alignment and shimming considerations involved. Further, splicing the raised portions 224 together could close off the ends of the stiffeners 214, thereby preventing sufficient water drainage and preventing visual inspection of any fasteners positioned under the raised portions 224.

Although the splice joint 106b of the illustrated embodiment is built up from a number of separate parts (e.g., the strap 220 and the fittings 230), in other embodiments, two or more of these parts can be integrated into a single part that performs the function and/or has the features of the two or more parts. For example, in one other embodiment, the splice joint 106b can be at least partially formed by a single part that integrates the features of the strap 220 and the fittings 230. In another embodiment, the splice joint 106b can include a single part that integrates the features of the strap 220 and the adjacent fillers 222. Although integrating parts may have the advantages of reducing part count and/or increasing strength, using separate parts may have the advantage of simplifying part construction and/or simplifying installation procedures.

FIGS. 3A-3C together illustrate a method of joining the first barrel section 104a to the second barrel section 104b in the vicinity of one of the window cutouts 140, in accordance with an embodiment of the invention. Referring first to FIG. 3A, this view is a partially exploded, enlarged isometric view looking outwardly at a portion of the second splice joint 106b around the window cutout 140. The portion of the first barrel section 104a illustrated in FIG. 3A includes a third panel portion 310a. The portion of the second barrel section 104b illustrated in FIG. 3A includes a fourth panel portion 310b positioned in edgewise alignment with the third panel portion 310a. The panel portions 310 can be at least generally similar in structure and function to the panel portions 210 described in detail above with reference to FIGS. 2A-2C. For example, the third panel portion 310a can include a plurality of stiffeners 214 (identified individually as stiffeners 214k-214m) attached to the first skin 112a. Similarly, the fourth panel portion 310b can include a plurality of stiffeners 214 (identified individually as stiffeners 214n-214p) attached to the second skin 112b. In one aspect of the illustrated embodiment, however, the window cutout 140 is formed in a third edge region 313a of the first skin 112a, and in an adjacent fourth edge region 313b of the second skin 112b.

Referring next to FIG. 3B, an elongate strap 320 is attached to the third edge region 313a of the first skin 112a and the adjacent fourth edge region 313b of the second skin 112b. With the exception of an aperture 324 that extends through a flared-out portion of the strap 320, the strap 320 can be at least generally similar in structure and function to the strap 220 described above with reference to FIGS. 2A-2C. For installation, the aperture 324 is aligned with the window cutout 140 and the strap 320 is attached to the skins 112 with a plurality of the fasteners 221. In other embodiments, the strap 320 can be bonded to the skins 112, or bonded and fastened to the skins 112.

One feature of the strap 320 is that the aperture 324 extends completely around the window cutout 140. One advantage of this feature is that the strap 320 acts as a one-piece doubler, thereby providing an efficient load path around the window cutout 140. A further advantage of this feature is that it reduces part count by combining the window doubler feature with the splice strap feature in a single, integrated part.

In the illustrated embodiment, the strap 320 is thicker than the adjacent flange portions 226 of the stiffeners 214. To avoid a step between adjacent surfaces, the first fillers 222a and the second fillers 222b are positioned on the flange portions 226 adjacent to the strap 320 in those portions of the splice joint 106b positioned away from the window cutout 140. Narrower fillers 322 (identified individually as third fillers 322a and fourth fillers 322b) are positioned on the stiffener flange portions 226 in those areas proximate to the window cutout 140.

Referring next to FIG. 3C, a plurality of the fittings 230 extend across the splice joint 106b in the stiffener bays away from the window cutout 140 as described above with reference to FIGS. 2A-2C. Narrower fittings 330 are attached across the splice joint 106b in similar fashion at opposing ends of the window cutout 140. The narrow fittings 330 of the illustrated embodiment have “L” shaped cross sections. In other embodiments, however, the narrower fittings 330 can have other cross sectional shapes, including “U” shapes, “C” shapes, and flat shapes. A window frame 350 can be fastened or otherwise attached to the strap 320 and any underlying structures around the window cutout 140. In one embodiment, the window frame 350 can be machined or otherwise formed from a high-strength metal material, such as aluminum. In other embodiments, the window frame 350 can include composites and/or other suitable materials.

One feature of the embodiments described above and illustrated in FIGS. 3A-3C is that the splice joint 106b extends through the middle of the window cutout 140. One advantage of this feature is that it provides design flexibility. For example, this feature allows window patterns and barrel section lengths to be selected irrespective of splice location. FIG. 4 is a cross-sectional end view of the splice joint 106b taken substantially along line 4-4 in FIG. 2C. This view illustrates that, in this embodiment, the fittings 230 are positioned over the strap 220, and the fasteners 238 extend through the fittings 230, the strap 220, and the skin 112b. This view further illustrates that the fittings 230 are positioned between, but proximate to, respective stiffeners 214.

The subject matter of co-pending U.S. patent application Ser. No. 10/646,509, entitled “MULTIPLE HEAD AUTOMATED COMPOSITE LAMINATING MACHINE FOR THE FABRICATION OF LARGE BARREL SECTION COMPONENTS,” filed Aug. 22, 2003; Ser. No. 10/717,030, entitled “METHOD OF TRANSFERRING LARGE UNCURED COMPOSITE LAMINATES,” filed Nov. 18, 2003; Ser. No. 10/646,392, entitled “AUTOMATED COMPOSITE LAY-UP TO AN INTERNAL FUSELAGE MANDREL,” filed Aug. 22, 2003; Ser. No. 10/630,594, entitled “COMPOSITE FUSELAGE MACHINE,” filed Jul. 28, 2003; Ser. No. 10/646,316, entitled “UNIDIRECTIONAL, MULTI-HEAD FIBER PLACEMENT,” filed Aug. 22, 2003; Ser. No. 10/301,949, entitled “PARALLEL CONFIGURATION COMPOSITE MATERIAL FABRICATOR,” filed Nov. 22, 2002; Ser. No. 10/799,306, entitled “SYSTEMS AND METHODS ENABLING AUTOMATED RETURN TO AND/OR REPAIR OF DEFECTS WITH A MATERIAL PLACEMENT MACHINE,” filed Mar. 12, 2004; Ser. No. 10/726,099, entitled “SYSTEMS AND METHODS FOR DETERMINING DEFECT CHARACTERISTICS OF A COMPOSITE STRUCTURE,” filed Dec. 2, 2003; Ser. No. 10/628,691, entitled “SYSTEMS AND METHODS FOR IDENTIFYING FOREIGN OBJECTS AND DEBRIS (FOD) AND DEFECTS DURING FABRICATION OF A COMPOSITE STRUCTURE,” filed Jul. 28, 2003; and Ser. No. 10/822,538, entitled “SYSTEMS AND METHODS FOR USING LIGHT TO INDICATE DEFECT LOCATIONS ON A COMPOSITE STRUCTURE, filed Apr. 12, 2004, is incorporated herein in its entirety by reference. In addition, the subject matter of U.S. Pat. No. 6,168,358 is also incorporated herein in its entirety by reference.

From the foregoing, it will be appreciated that specific embodiments of the invention have been described herein for purposes of illustration, but that various modifications may be made without deviating from the spirit and scope of the invention. For example, aspects described in the context of particular vehicles, such as aircraft, can equally apply to other vehicles, such as helicopters, rockets, watercraft, etc. Further, aspects described in the context of particular embodiments can be combined or eliminated in other embodiments. Accordingly, the invention is not limited, except as by the appended claims.

The description of the different advantageous embodiments has been presented for purposes of illustration and description, and is not intended to be exhaustive or limited to the embodiments in the form disclosed. Many modifications and variations will be apparent to those of ordinary skill in the art. Further, different advantageous embodiments may provide different advantages as compared to other advantageous embodiments. The embodiment or embodiments selected are chosen and described in order to best explain the principles of the embodiments, the practical application, and to enable others of ordinary skill in the art to understand the disclosure for various embodiments with various modifications as are suited to the particular use contemplated.